The present invention relates generally to arrangements for launching payloads, such as spaceraft or satellites from launch vehicles, and more particularly, to launching arrangements in which the payload is mechanically restrained prior to its gyroscopic ejection from the launch vehicle in order to dissipate unwanted forces.
A publication entitled "Syncom IV Space Shuttle Orbital Flight Test Mission," publication number SCG 66710V/Dec. 1976 describes an arrangement for launching a spacecraft from a launch vehicle. The arrangement comprises a U-shaped or open-ended cradle having an ejection spring mechanism located on one side thereof for pushing against a small trunion that protrudes from one side of the spacecraft. A pivot point is formed on the opposite side of the spacecraft which also has a protruding trunion that rests on a mating surface formed in the cradle.
Although not described in this publication, it is necessary that both the spring mechanism and the pivot point lie in a plane normal to the spin axis of the spacecraft. Ideally, the plane also passes through the center of mass of the spacecraft. Release of the ejection spring mechanism applies a tangential thrust force, acting in the plane, to the spacecraft. Assuming the geometry of the structure is precise, when the spacecraft tangential thrust force is applied, the spacecraft rotates about the pivot point, simultaneously producing translational and rotational movement of the spacecraft without nutation as it leaves the cradle.
After the ejection force ceases, the spacecraft free body motion is a rolling motion up an imaginary ramp, thus maintaining the imparted linear and angular momentum. The separation velocity and rotational speed of the spacecraft depend on its inertia characteristics, diametral dimensions, ejection force and ejection stroke length.
The payload deployment system described in U.S. Pat. No. 4,359,201 entitled "Payload Deployment From Shuttle with Linear and Angular Velocity," assigned to the assignee of the present invention is an improvement over the system described in the above-cited publication. The deployment system described in this patent utilizes a single point ejection force for simple payload deployment mechanization. The force application and spacecraft/cradle reaction points enclose the payload center of mass and center of percussion. This minimizes the need for elaborate hold-down mechanisms at the force reaction points and for precise alignment of the push-off force with the center of mass and the reaction pivot point.
Structurally this involves, at a minimum, a pair of spaced pivot points on one side of the spacecraft and a force application point on the other side supporting the spacecraft in the open end of the cradle. These points define a triangle which encloses the center of mass of the spacecraft. Such a three-point suspension system defines the spacecraft attitude at separation from the launch vehicle. Rotation of the payload or spacecraft about fixed supports, such as the spaced pivot points, ensures physical clearance during the ejection phase of the launch as well as a well-defined deployment path. Also, the affects of spacecraft attitude disturbances due to liquid propellant sloshing are avoided while the spacecraft is being separated and in physical contact with the shuttle or launch vehicle.
Although the deployment system described in this patent has provided a significant improvement in payload launching systems, this design requires rather stiff structural supports, since the pivot points must support the payload structure. There is a possibility that thermal and gravity loads associated distortions of these stiff structures may cause a phenomenon known as "gapping" at the pivot points.
In addition, the phenomenon of "racking" may occur due to potentially large, statically indeterminant forces exerted on the payload at the moment of deployment by the stiff support structure. The use of a rigid support structure acts to compound the gapping and racking problems and also tends to cause "pivot bouncing" at the moment of deployment. These phenomena cause undesirable rotational effects which affect the attitude of the payload.
In typical deployment situations, several release mechanisms securing the spacecraft to the cradle are activated, after which forces were exerted on the spacecraft by the spring-loaded force actuator. The activation of these release mechanisms also created undesirable forces and moments to be exerted upon the spacecraft. It is therefore beneficial to eliminate these potential problems in order to have a well-controlled payload launch.